1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine blade with leading edge cooling.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, a combustor produces a extremely high temperature gas flow that is passed through a turbine to produce mechanical power. The turbine typically includes multiple stages or stator guide vanes and rotor blades that are exposed to the hot gas flow. The first stage stator vanes and rotor blades are exposed to the highest temperature, since the temperature progressively decreases as the hot gas flow passes through the turbine stages and energy is extracted from the flow.
It is well known in the art of gas turbine engines that the engine efficiency can be increased by increasing the inlet temperature of the hot gas flow into the turbine. A higher gas flow temperature means higher energy content in the flow. However, the limiting temperature entering the turbine first stage is dependent upon the material characteristics of the vanes and blades as well as the cooling capability. Thus, the turbine inlet temperature can be increased with improved materials and/or improved cooling.
A turbine blade or vane will be exposed to different levels of temperature throughout the airfoil surface. Complex internal cooling passages are designed so that adequate cooling of each surface of the airfoil can be accomplished. In an area where too little cooling is produced, a hot spot can occur in which erosion or oxidation of the airfoil surface will occur and lead to damaged airfoils. Especially in an industrial gas turbine engine, where the engine can operate continuously for 48,000 hours, a damaged airfoil may result in premature stopping of the engine for repairs or a damaged airfoil that will result in lower performance of the engine. Thus, to provide for long engine runs and long part life, the turbine vanes and blades require maximum cooling over all the surfaces and minimal amounts of cooling air to provide for an increased efficiency of the engine.
The highest heat load applied to the stator vanes and the rotor blades occur on the leading edge of the airfoil, since this area of the airfoil is exposed directly head-on to the hot gas flow. In the prior art, a blade leading edge showerhead comprises three rows of film cooling holes. The middle film row is positioned at the airfoil stagnation point where the highest heat load occurs on the airfoil leading edge. Film cooling holes for each film row are at an inline pattern and inclined at 20 to 35 degrees relative to the blade leading edge radial surface. FIGS. 1 and 2 show a cutaway view of the prior art blade leading edge showerhead film cooling hole arrangement. A fundamental shortfall associated with this prior art showerhead arrangement is the over-lapping of film ejection flow in a rotational environment such as a rotor blade. Stator vanes do not rotate, and therefore this problem is not the same as with rotor blades because of the centrifugal forces associated with the cooling air ejection from the film cooling holes. As a result of the prior art film hole arrangement, a hot streak is developed as shown in-between the film holes of FIG. 4 which shows the film trace for the turbine blade showerhead of the prior art.
One prior art reference, U.S. Pat. No. 7,114,923 B2 issued to Liang on Oct. 3, 2006 and entitled COOLING SYSTEM FOR A SHOWERHEAD OF A TURBINE BLADE, discloses a showerhead arrangement in which the film cooling holes are extending at various angles relative to each other in order to reduce the likelihood of zipper effect cracking in the leading edge and to effectively cool the leading edge of the turbine blade. The Liang U.S. Pat. No. 7,114,923 is incorporated herein by reference in its entirety.
It is therefore an object of the present invention to provide for a turbine blade with a showerhead arrangement of film cooling holes that will produce a more effective film cooling covering of the leading edge than the cited prior art references.